A gas turbine engine, typically used as a source of propulsion in aircraft, operates by drawing in ambient air, mixing and combusting that air with a fuel, and then directing the exhaust from the combustion process out of the engine. A compressor having a low-pressure and high-pressure compressor, rotates to draw in and compress the ambient air. A portion of the compressed air is used to cool a combustor, while the rest is directed into the combustor, mixed with a fuel, and ignited.
Typically, an igniter generates an electrical spark to ignite the air-fuel mixture. The products of the combustion, NOx and CO, then travel out of the combustor as exhaust and through a turbine. The turbine, having a low-pressure and high-pressure turbine, is forced to rotate as the exhaust passes through the turbine blades. The turbine and the compressor are connected by concentrically mounted rotating shafts running through the center of the engine, one shaft for the low-pressure compressor and turbine and one shaft for the high-pressure compressor and turbine. Thus, as the turbine rotates from the exhaust, the compressor rotates to bring in and compress new air. Once started, it can therefore be seen that this process is self-sustaining.
Combustors for gas turbine engines typically have an outer shell and an outer liner, disposed radially inside the outer shell. Additionally, annular combustors have an inner shell and an inner liner radially outside the inner shell. The inner and outer liners are separated by and define a combustion chamber. Flow cavities are typically provided between each pair of shells and liners. Cooling air is forced through these flow cavities and into the combustion chamber, creating a cooling film on hot surfaces of the liners.
Prior art combustion chamber configurations used geometrical profiles that were not convergent in a primary rich zone. Over time, however, such configurations evolved to have a convergent section at the primary rich zone. One of the design intents of such prior art designs was to increase the combustion flow velocity to reduce corresponding combustor residence time. Since time plays a direct part in NOx formation, convergent combustion chamber designs provided an added benefit for NOx. However, aggressive tapering of the convergent combustion chamber section may cause entrainment of cooling flow in the outer recirculation zone. This, in turn, may effect local chemistry as the fuel rich zone trends towards stoichiometric conditions. In this case, flame temperatures increase along with NOx formation.
In light of the foregoing, it can be seen that the gas turbine engine industry continues to strive for designs with reduced NOx emissions, while at the same time increasing engine cooling to thus enhance the serviceable life of the engine.